Axial blade slot pressure face with undercut

ABSTRACT

A blade and disk assembly includes a blade body and a disk body having a slot to receive the blade body. The blade body has a blade bearing surface. The slot has a disk bearing surface that contacts the blade bearing surface at a bearing interface. The slot has an undercut surface that is spaced apart from the blade body. An angle between the disk bearing surface and a line tangent to the undercut surface is at least 70 degrees.

BACKGROUND OF THE INVENTION

This disclosure relates to a blade and disk interface and, more particularly, to forming an undercut at the blade and disk interface to reduce stresses.

Gas turbine engine components, such as turbine blades, turbine vanes, compressor blades, compressor vanes, or other components typically operate in a relatively high stress and high temperature environment. The stresses and temperature may result in damage to the component from corrosion, erosion, deformation, or the like.

One example of a high stress area is an interface between a blade component and an associated disk component that supports the blade component for rotation about an engine axis. For example, a disk body for a turbine or compressor includes a plurality of slots with one blade component being secured within each slot. Typically, the blades have a dove-tail formation at one end that is slid into the slot from one side of the disk. The dove-tail formation includes a blade bearing surface that contacts a disk bearing surface formed as part of a surface that defines the slot. When the gas turbine engine is operation, the blades slide radially outward within the slot due to centrifugal loading. This relative movement generates stresses at certain interfaces between the blade and the disk, such as crush stresses, shear stresses, hoop stresses, etc.

One proposed solution utilizes an undercut in the slots to reduce stresses by minimizing contact area between the blade and disk at high-stress contact areas. A surface of the slot defines an angle that is formed between a line extending along a blade and disk bearing interface and a line tangent to the undercut. In one example, this angle is 14 degrees; however, it has been known to have this angle approach up to 60 degrees.

Traditionally, angles between 3 to 60 degrees have been recommended as providing the most significant stress reduction. Angles in excess of 90 degrees are disfavored because such high angles weaken the disk. This can cause cracking, which can result in pre-mature wear or failure. However, as gas turbine engines are configured to operate under higher speeds and more severe conditions, stresses at the blade and disk interface continue to be analyzed for further stress reduction opportunities.

SUMMARY OF THE INVENTION

An example blade and disk assembly includes a blade body and a disk body having a slot to receive the blade body. The blade body has a blade bearing surface. The slot has a disk bearing surface that contacts the blade bearing surface at a bearing interface. The slot has an undercut surface that is spaced apart from the blade body. An angle between the disk bearing surface and a line tangent to the undercut surface is at least 70 degrees.

In another aspect, the undercut surface transitions directly from the disk bearing surface to form an undercut in the slot.

In another aspect, the line tangent to the undercut surface is defined at a location where the undercut surface transitions from the disk bearing surface.

In one example, the angle is 70 degrees or greater and is less than 90 degrees.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiment. The drawings that accompany the detailed description can be briefly described as follows.

FIG. 1 is a schematic view of a gas turbine engine.

FIG. 2A is a perspective view of a disk as used in the gas turbine engine of FIG. 1.

FIG. 2B is an enlarged view of a portion of the disk of FIG. 2A.

FIG. 3 is a schematic cross-sectional view of a blade and disk interface.

FIG. 4 is a detail view of an angle defined at the blade and disk interface of FIG. 3.

FIG. 5 is a chart detailing percentage stress reductions when compared to prior configurations.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 illustrates selected portions of an example turbine engine 10, such as a gas turbine engine 10 used for propulsion. In this example, the gas turbine engine 10 is circumferentially disposed about an engine centerline 12. The turbine engine 10 includes a fan section 14, a compressor section 16, a combustion section 18, and a turbine section 20. The compressor section 16 and the turbine section 20 include corresponding blades 22 and vanes 24. The turbine section 20 includes a high pressure turbine (HPT) section 20 a and a lower pressure turbine (LPT) section 20 b. Further, the compressor section 16 includes a high pressure compressor (HPC) section 16 a and a low pressure compressor (LPC) section 16 b. As is known, air compressed in the compressor section 16 is mixed with fuel and burned in the combustion section 18 to produce hot gasses that are expanded in the turbine section 20.

FIG. 1 is a schematic presentation for illustrative purposes only and is not a limitation on the disclosed examples. Additionally, there are various types of gas turbine engines, many of which could utilize the examples disclosed herein and are not limited to the designs shown.

FIG. 2 illustrates an example of a disk section 30 to which the blades 22 are attached. The disk section 30 includes a disk body 32 having a plurality of slots 34. Each slot 34 receives one blade 22 as shown in FIG. 3. The slots 34 define pressure faces relative to the blade 22. A concave pressure face 36 comprises a left side pressure face as shown in FIG. 2B, and a convex pressure face 38 comprises a right side pressure face. The disk section 30 and associated blades 22 rotate together about the engine centerline 12.

FIG. 3 shows a portion of the disk body 32 with one end of the blade 22 received within the slot 34. The blade 22 has a blade body 36 with a dove-tailed end 38 that is slid from the side of the disk body 32 into the slot 34 as known. The blade body 36 includes an airfoil portion at an end opposite from the dove-tailed end 38. The air foil portion is positioned to rotate relative to the vanes 24 as known.

The blade body 36 includes a blade bearing surface 40 that extends at an oblique angle to a surface 42 that extends outward of the slot 34. The blade bearing surface 40 comprises a linear surface that faces a surface that defines the slot 34.

The slot 34 is defined by a disk bearing surface 50 that is positioned opposite from the blade bearing surface 40. The disk bearing surface 50 comprises a linear surface that contacts the blade bearing surface 40 at a bearing interface during engine operation. An exaggerated gap is shown at the bearing interface for clarity purposes.

It is known that a high stress concentration area is formed at an interface between a transition edge 52 of the blade body 36 and a corresponding surface of the slot 34. In order to reduce stresses at this contact interface, an undercut 60 has been formed within the slot 34. The undercut 60 comprises a curved surface and transitions directly from the disk bearing surface 50 at each edge. As such, the undercut surface is spaced apart from the blade body 36 by a gap.

In this configuration, an angle A is defined between a line L1 extending along the bearing interface as the blade bearing surface 40 contacts the disk bearing surface 50, and a line L2 that is tangent to the undercut 60 at a point where the disk bearing surface 50 transitions to the undercut surface. This is shown in greater detail in FIG. 4.

In one example, this angle A is at least 70 degrees. FIG. 4 shows a profile of the undercut 60 and the disk bearing surface. In this example, the line L1 and the line L2 are orientated at an angle A that is 70 degrees.

Traditionally, angles larger than 60 degrees have been disfavored due to potential weakening of the disk body 32 at the undercut location. However, analysis has been performed, and it has been found that increasing the angle A to 70 degrees provides significant reductions in stresses.

FIG. 5 is a chart that shows comparisons between a traditional (no-undercut) configuration, a 14 degree configuration, and a 70 degree configuration. For each configuration the following stresses were measured: 1) maximum principal stresses MPS1 at corresponding convex 38 and concave 36 pressure faces; 2) minimum principal stresses MPS2 at corresponding convex 38 and concave 36 pressure faces; and 3) hoop stresses HS3. The analysis was performed for a bladed disk without friction at a maximum stress time point. The analysis showed that the 70 degree configuration had lower stresses than the traditional and 14 degree configurations.

As shown in FIG. 5, when a comparison is made between the 14 degree and 70 degree configurations, the maximum principal stresses MPS1 were reduced at the convex surface by 8.2% and at the concave surface by 1%. The minimum principal stresses MPS2 for the 70 degree configuration were reduced at the convex surface by 25.3% and at the concave surface by 6% when compared to corresponding stresses on the 14 degree configuration. Further, hoop stresses HS3 were reduced by 0.1%. Thus, as shown, the 70 degree configuration provides significant stress reduction compared to prior designs without weakening the disc.

Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments.

The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims. 

1. A blade and disk assembly comprising: a blade body having a blade bearing surface; a disk body including at least one slot to receive said blade body, said slot having a disk bearing surface that contacts said blade bearing surface at a bearing interface, and said slot having an undercut surface that is spaced apart from said blade body, and wherein an angle between said disk bearing surface and a line tangent to said undercut surface is at least 70 degrees.
 2. The blade and disk assembly according to claim 1 wherein said angle is less than 90 degrees.
 3. The blade and disk assembly according to claim 1 wherein said undercut surface transitions directly from said disk bearing surface to form an undercut in said slot.
 4. The blade and disk assembly according to claim 3 wherein said undercut surface comprises a curved surface and said disk bearing surface comprises a linear surface.
 5. The blade and disk assembly according to claim 4 wherein said blade bearing surface comprises a linear surface.
 6. The blade and disk assembly according to claim 3 wherein said line tangent to said undercut surface is defined at a location where said undercut surface transitions from said disk bearing surface.
 7. A gas turbine engine comprising: a fan; a compressor section; a combustor; and a turbine section wherein said compressor and said turbine sections each include a disk body having a plurality of slots with each slot receiving one blade, and wherein at least one of said plurality of slots includes a first bearing surface that contacts a second bearing surface on a corresponding blade at a bearing interface, said slot having an undercut surface that is spaced apart from said corresponding blade, and wherein an angle between said first bearing surface and a line tangent to said undercut surface is at least 70 degrees.
 8. The gas turbine engine according to claim 7 wherein said undercut surface transitions directly from said first bearing surface to form an undercut in said slot.
 9. The gas turbine engine according to claim 8 wherein said line tangent to said undercut surface is defined at a location where said undercut surface transitions from said first bearing surface.
 10. The gas turbine engine according to claim 9 wherein said undercut surface comprises a curved surface and said first bearing surface comprises a linear surface.
 11. The gas turbine engine according to claim 10 wherein said second bearing surface comprises a linear surface.
 12. The gas turbine engine according to claim 7 wherein said angle is less than 90 degrees. 